1. Field of the Invention
The invention relates to the field of gas turbine engines, and more specifically, to an improved turbine rotor and a gas turbine engine using compressor fluid to maintain a thermal boundary layer between turbine blades and heated fluid from the combustor.
2. Description of Related Art
A type of prior art gas turbine has a compressor, a fuel source, a source of air for combustion, a casing, and a combustor. The combustor has a combustion zone that is connected to a fuel source and a source for combustion air. It contains a cooling zone for cooling the resulting heated fluid before reaching the turbine. The combustor cooling zone is connected to the compressor. The heated fluid temperature fluctuates depending on operating conditions. In conventional gas turbine engines, these temperature fluctuations result in strong temperature-induced stresses imparted on the engine components.
In these prior art gas turbines, virtually the entire compressor fluid flow is directed to the combustor. Fluid heated in the combustor is cooled by the compressor fluid flow in the combustor cooling zone. This engine has a turbine rotor disk with blades that receive heated fluid from the combustor. The temperature of this heated fluid is quite high and, under certain conditions, the fluid can overheat the turbine rotor disk blades. To prevent such overheating, each blade has an interior channel that receives air from the compressor. As a result, part of the fluid coming from the compressor (about 3% to 5% of the total flow) is supplied to the interior channel of the blades to keep their temperature within design limits. One example of this gas turbine engine is disclosed in U.S. Pat. No. 3,826,084 to Branstrom et al.
The heated fluid also has to be cooled after fuel combustion. Normally, this is done in the combustor, to which the major part of the fluid from the compressor would be admitted.
Thus, in this type of prior art gas turbine, substantially all of the fluid coming from the compressor is supplied to the combustor cooling zone to cool the fluid before it enters the turbine. When fluid from the compressor is mixed with the heated fluid in the combustor cooling zone, about 3% to 5% of the fluid""s energy is lost. Diverting about 3% of fluid from the compressor to the turbine rotor disk blades results in another 3% in losses. In addition, the combustor for this prior art gas turbine has to be made larger to accommodate the cooling zone.
It is also known to operate a gas turbine engine having a compressor for producing a compressed fluid flow, a casing, a combustion zone in the casing, a power turbine rotor disk with blades, each having an external airfoil surface with a leading surface and a trailing surface, an inlet edge positioned immediately downstream of the combustion zone, and an outlet edge positioned downstream of the inlet edge, a blade flow portion located adjacent said trailing surface of said external airfoil surface and between said inlet edge and the outlet edge, wherein the compressed fluid is supplied from the compressor to the blade flow portion, fuel and air for combustion are supplied to the combustion zone to prepare a heated fluid, and the heated fluid from the combustion is supplied directly to the blade flow portion (e.g., U.S. Pat. No. 6,305,157 to Rakhmailov). In this method, compressed fluid flow from the compressor supplied to the blade flow portion amounts to between 55% and 85% of the total fluid flow from the compressor. The fluid from the compressor forms a protective layer on the blade flow portion of the blades, preventing the hot fluid from the combustion zone from coming into direct contact with the blade material. Since the fluid from the compressor that was being fed to the combustor dilution zone for cooling down the hot fluid before supplying it to the turbine blade is now used for blade protection, losses in the combustor dilution zone are eliminated, and the fluid directed from the compressor to the blades performs work of expansion by adding to the energy.
A gas turbine engine for carrying out this method has a power turbine rotor disk with blades, each having an internal passage for receiving the compressed fluid from the compressor and passing the compressed fluid through at least one opening to the blade flow portion. The compressed fluid flow moves along the blade flow portion to create a thermal insulating boundary between the heated fluid and the trailing surface of the external airfoil surface.
This type of gas turbine engine has an internal passage in the blade, which make it harder to manufacture the blades. It is known that providing cooling passages in conventional gas turbine blades has always been a problem, making the manufacture of the blades more difficult. Moreover, the presence of internal passages and openings in the blades compromise their strength and durability and requires additional measures to be taken to assure reliability of blades in operation. All this makes this type of gas turbine blade more expensive and labor intensive to manufacture, and reduces the service life of the blades. In addition, the flow of the fluid from the compressor to the blade flow portion in using the above-described method of providing an insulating layer over the blade surface is restricted because the flow has to pass through slits in the blade body, which causes losses. Further, these slits in the blade body cannot be positioned in certain areas of the blade, making it difficult to assure uniform distribution of the fluid flow over the blade surface.
It is an object of the invention to provide a turbine rotor for a gas turbine engine that is easier to manufacture.
It is another object of the invention to assure better distribution of the compressed fluid flow over the blade surface with lower drag losses.
Another object of the invention is to provide a gas turbine rotor having blades that do not have internal passages.
A further object of the invention is to prolong the life of the gas turbine engine.
The foregoing and other objects are accomplished through the design of a gas turbine engine rotor having a body and blades, which are attached to the rotor by a root portion, for receiving a hot fluid flow and a compressed fluid from a compressor. Each blade has a base portion, an external airfoil surface defined by a blade body with an inlet edge positioned downstream of the combustion zone, an outlet edge positioned downstream of the inlet edge, a blade flow portion of the airfoil located between the inlet edge and said outlet edge on either side of the blade, and a passage system communicating with the compressor and with the blade flow portion to create a thermal insulating boundary between the heated fluid and the external airfoil surface. The rotor has a substantially planar platform member, extending substantially transversally to the body of the blade and having an upstream portion protruding beyond the inlet edge in the direction toward the combustion zone. The platform member has at least a pair of through openings on either side of the airfoil, said openings being positioned in series along said blade flow portion. One opening of the pair that is remote from the inlet edge defines an inclined passage in the platform member, and the other opening is separated by a partition from the first opening and has an opposite wall extending in a spaced relation to the inlet edge.
With this construction, the body of the blade does not have any internal passages or slits, and all flows, including the compressed fluid flow coming from the compressor for cooling the blade by insulating it from the hot fluid flow, occur outside the blade body. This facilitates manufacture of the blade and also improves strength and durability of the blade, prolonging the blade life and enhancing its reliability.
The airfoil surface of the blade has a guide portion, which extends beyond the platform member toward the root portion, and the platform member and the guide portion define nozzles on either side of the airfoil surface to increase velocity of the compressed fluid flow.
Other objects and advantages of the invention will become apparent from the following detailed description of preferred embodiments and accompanying drawings.